CN106742075B - Distributed propulsion system - Google Patents
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- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
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- B64G1/40—Arrangements or adaptations of propulsion systems
Abstract
The input end of a gas collecting device of a high-energy working medium collecting device is communicated with the output end of a core machine compressor of a turbine engine core machine; the output end of the high-energy working medium collecting device is communicated with the inlet of the high-efficiency working medium transmission device. The output end of a transmission branch pipe in the high-efficiency working medium transmission device is respectively communicated with the input end of a propeller turbine volute of each distributed propeller; the input end of the transmission branch pipe is communicated with a transmission main pipe in the high-efficiency working medium transmission device. 6 distributed propellers are evenly distributed on two sides of the core machine of the turbine engine. The invention overcomes the dependence of the distributed propulsion system based on power distribution on high-energy density energy storage equipment and a super-high power motor, and improves the realizability of the distributed propulsion system; meanwhile, the mechanical constraint between the gas generator and the propeller in the mechanical transmission-based distributed propulsion system is eliminated, the bypass ratio limit of the turbofan engine and the traditional distributed power is broken through, the regenerative design is realized, and the propulsion efficiency is improved.
Description
Technical Field
The invention relates to an aviation propulsion system, in particular to a distributed propulsion system.
Background
The economy is a key factor to be considered by the conveyor, and along with the continuous highlighting of environmental problems, people also put more and more rigorous requirements on reducing carbon emission. This all requires an increase in the efficiency of the power of the existing transport (turbofan engine). One of the effective measures to improve the efficiency of turbofan engines is to use a higher bypass ratio to achieve high propulsion efficiency. The bypass ratio of the prior civil turbofan engine is close to 10, but the bypass ratio is difficult to be further improved due to the limitation of structure and component matching. The distributed propulsion system breaks through the structural and component matching limitations of the traditional turbofan engine, and therefore higher bypass ratio can be obtained. Meanwhile, the distributed propulsion system is also more suitable for serving as power of a future wing body fusion aircraft.
United kingdom roclo company has obtained in 2012-2013 a number of patents on the invention of "distributed propulsion system and control method thereof", patent numbers: EP2581308A2, US2013/0094963 A1, US9376213B2. The invention adopts two turbine engines arranged under the wing to drive the engine to generate power, then the power is transmitted to the propellers distributed on the wing, the wing tip or two sides of the rear part of the fuselage, and the motor drives the propellers to generate thrust. The rolo corporation also applied for a similar chinese patent in 2014, patent number: CN 104670503A.
United states technology corporation obtained an invention patent named "counter-rotating rotor distributed propulsion system" in 2015, patent No.: EP2930114A1, US 2015/0284071 A1. The idea of the invention is that a gas generator positioned at the tail part of a machine body is used for driving a power turbine to rotate, and the power turbine indirectly drives counter-rotating rotor-rotating rotors positioned at two sides of a rear machine body through a main speed reducer and a secondary speed reducer to generate thrust.
United states combined technology corporation proposed an aircraft propulsion system in the invention creation publication US 2008/0098719 A1. In this system, a single gas generator drives a low-pressure turbine which indirectly drives a plurality of fans on both sides of the fuselage through a two-stage reducer. The air flow compressed by the fan is divided into two parts, one part is discharged from the tail of the aircraft through a duct, and the other part is sucked by the compressor and participates in the thermodynamic cycle of the gas generator.
The air passenger company discloses in publication number CN 104229144A the invention creation of an aircraft with an electrical device. The invention is essentially an electric energy based distributed propulsion system. The electric energy is generated by the electric energy generator, and then the electric energy is distributed to the propelling devices positioned on the two sides of the fuselage through the power supply device so as to drive the propelling devices to generate thrust. In order to solve the problem that the output power of the electric energy generator is not matched with the required power of the propeller, the system is also provided with an energy storage device and a hybrid power system.
The invention and creation with the publication number CN 104973234A provides an aircraft adopting a distributed electric ducted fan flap lift-rising system. The aircraft is characterized in that a plurality of ducted fans positioned on wings and a lift fan system positioned at the rear part of an aircraft body are driven by a power source. It is not indicated which power source is used.
There is no academic paper or academic paper about distributed propulsion systems.
The essence of a distributed propulsion system is to distribute the energy generated by a centralized energy source generator to a plurality of distributed propellers, with the distribution and transmission of energy. Existing distributed propulsion systems can be classified into two categories according to the energy distribution transmission mode: one is based on electric drive (patents: EP2581308A2, US2013/0094963 A1, US9376213B2, CN 104670503A, CN 104229144A, CN 104973234A) and the other is based on mechanical drive (patents: EP2930114A1, US 2015/0284071 A1, US 2008/0098719 A1). In both systems, the high temperature and high pressure combustion gases generated by the gas turbine engine are used to generate electricity or to drive a turbine to generate shaft work, and then the electricity and the shaft work are used to drive a propeller to operate to generate thrust. Both systems still present some technical difficulties in implementation. The technical challenge of distributed propulsion systems based on power distribution is the development of high density energy storage devices and super high power motors. The problem of the distributed propulsion system based on mechanical transmission lies in that it does not completely get rid of the mechanical constraints of the gas generator and the propellers, is limited by the factors such as the structure and weight of the mechanical transmission device, and the propellers are distributed at a limited distance, which is not favorable for the layout on the aircraft.
Disclosure of Invention
The power distribution system aims to overcome the defect that the power of the existing energy storage equipment and the motor cannot meet the requirement in the prior art based on power distribution; based on the defects that a mechanical transmission mechanism is complex and heavy and is not beneficial to arrangement on an aircraft during mechanical transmission, the invention provides a distributed propulsion system.
The invention comprises a turbine engine core machine, a high-efficiency working medium transmission device, a high-energy working medium acquisition device and a distributed propeller. The input end of the gas collecting device of the high-energy working medium collecting device is communicated with the output end of a core machine compressor of the turbine engine core machine; and the output end of the high-energy working medium collecting device is communicated with the inlet of the high-efficiency working medium transmission device. The output end of a transmission branch pipe in the high-efficiency working medium transmission device is respectively communicated with the input end of a propeller turbine volute of each distributed propeller; the input end of the transmission branch pipe is communicated with a transmission main pipe in the high-efficiency working medium transmission device. 6 distributed propellers are uniformly distributed on two sides of the turbine engine core machine.
The turbine engine core is a single-rotor gas turbine jet engine or a double-rotor turbine jet engine; the distributed propellers are ducted fans driven by propeller turbines or propellers driven by the propeller turbines.
The gas collecting device comprises four gas guide tubes, a gas collecting ring and four gas guide tube regulating valves. The gas collecting ring shell is uniformly provided with four gas guide pipe connecting holes on the surface of the same side, and the input ends of the four gas guide pipes are respectively arranged on the gas guide pipe connecting holes. The four air duct regulating valves are respectively arranged at the output ends of the air ducts. The surface of the gas collecting ring shell is arc-shaped, and the opening of the arc-shaped gas collecting ring shell is positioned on the inner side of the gas collecting ring shell, so that the cross section of the arc-shaped gas collecting ring shell is U-shaped. Cutting the compressor outer casing into two sections; and respectively and fixedly connecting two side walls of the opening of the gas collecting ring with the compressor outer casing which is divided into two sections, so that the gas collecting ring and the compressor outer casing jointly form an outer ring channel of the gas flow at the outlet of the compressor. The core compressor on the gas turbine engine core machine is provided with two annular outlets, namely an inner annular outlet and an outer annular outlet; the outer annular outlet is communicated with the input end of the outer annular channel of the air compressor outlet airflow; the inner annular outlet is communicated with the input end of the core engine combustion chamber.
During assembly, the output ends of the four gas guide pipes on the gas collecting ring shell are respectively communicated with the input ends of four tail gas heat recovery devices distributed in a rear casing of the turbine, and the gas flow of an outer ring channel of the gas flow at the outlet of the gas compressor is transmitted into each tail gas heat recovery device through each gas guide pipe. The 4 tail gas heat recovery devices are divided into two groups, and the output ends of the tail gas heat recovery devices of each group are respectively communicated with the transmission main pipes positioned on two sides of the core machine of the turbine engine through hot air guide pipes.
The tail gas heat recovery device is circumferentially and uniformly arranged on the inner side of the turbine rear casing, and an included angle alpha between the axis of the tail gas heat recovery device and the axis of the turbine engine core machine is = 0-90 degrees.
The two transmission main pipes are respectively positioned on two sides of the turbine engine core. The output ends of the U-shaped tube bundle heat regenerators are communicated with the hot air guide tubes and are respectively communicated with the input ends of the transmission main pipes through the three-way joints. Three airflow output ends are distributed on each transmission main pipe, and each airflow output end is connected with a transmission branch pipe. And the output end of the transmission branched pipe is respectively communicated with the input end of the propeller turbine volute of each distributed propeller.
And the input end of the transmission branch pipe penetrates through the through hole on the distributed propeller support plate and is connected with the input end of the propeller turbine volute. The propeller turbine volute is annular and is wrapped around the propeller turbine, and the output end of the propeller turbine volute is uniformly distributed on the inner side of the propeller turbine volute and is connected with the input end of the centrifugal propeller turbine.
When the distributed propeller is a ducted fan driven by a propeller turbine, the ducted fan comprises a fan, a ducted fan casing, a gear reducer, a propeller turbine shaft, a propeller turbine volute, a propeller turbine, a distributed propeller inner casing and a distributed propeller support plate. The 4 distributed propeller support plates are uniformly distributed between the ducted fan casing and the distributed propeller inner casing, so that one end of each distributed propeller support plate is fixed on the inner surface of the ducted fan casing, and the other end of each distributed propeller support plate is fixed on the outer surface of the distributed propeller inner casing. The propeller turbine shaft is positioned in the inner casing of the distributed propeller, and two ends of the propeller turbine shaft are both arranged on the inner support of the distributed propeller through bearings; the inner support of the support distributed propeller is fixed on the inner surface of a casing in the distributed propeller; the center line of the propeller turbine shaft is superposed with the center line of the distributed propeller inner casing.
When the gas turbine engine core machine adopts a dual-rotor turbojet engine, the output end of a high-pressure compressor of the gas turbine engine core machine is communicated with the input end of a gas collecting device of a high-energy working medium collecting device; and the output end of the high-energy working medium collecting device is communicated with the inlet of the high-efficiency working medium conveying device.
When the distributed thruster is a propeller driven by a thruster turbine, the propeller pushes the airflow to move backwards to generate pulling force.
The essential difference between the distributed propulsion system and the existing distributed propulsion system is that the distribution and transmission of energy in the distributed propulsion system are realized by adopting a working medium transmission mode, namely, high-energy working medium generated by an energy source is directly transmitted into the distributed propellers to drive the propellers to operate. On one hand, the distributed propulsion system based on working medium transmission can overcome the dependence of the distributed propulsion system based on electric power distribution on high-energy-density energy storage equipment and an ultra-high-power motor, and the realizability of the distributed propulsion system is improved; on the other hand, the mechanical constraint between the gas generator and the propeller in the mechanical transmission-based distributed propulsion system can be released, the bypass ratio limit of the traditional turbofan engine and the traditional distributed power is broken through, the regenerative design is realized, and the propulsion efficiency is improved.
The invention uses the prior volute technology and U-shaped tube bundle heat regenerator technology. The volute is a hollow shell wrapped outside the centripetal turbine, and guide vanes are arranged in the volute and used for collecting external airflow and guiding the external airflow into the centripetal turbine to push the centripetal turbine to do work. Volutes are a common structure for aircraft turbine engines and are also widely used in turbochargers.
The heat regenerator is a core component in the intercooling heat-regenerative turbine engine, and the air flow at the output end of the air compressor and the gas at the output end of the turbine exchange heat through the heat regenerator, so that the waste heat of the gas at the output end of the turbine is effectively utilized, and the working capacity of the air flow at the output end of the air compressor is improved. The U-shaped tube bundle heat regenerator is one of heat regenerators used for intercooling heat regeneration turbine engines, and has the main structure that a plurality of groups of U-shaped tubes are arranged on two sides of a gas collecting tube and a gas collecting tube, heat exchange airflow enters from the gas inlet gas collecting tube and flows into the gas outlet gas collecting tube through two opposite U-shaped tubes, and gas at the output end of a hotter turbine sweeps across a heat exchanger at a certain incoming flow attack angle, so that the purpose of heat exchange is achieved.
The distributed propulsion system based on working medium transmission comprises a gas turbine engine core machine, a high-energy working medium collecting device, a high-efficiency working medium transmission device and a distributed propeller. The working mode is that the gas turbine engine core machine generates high-energy working medium through internal thermodynamic cycle. The high-energy working medium collecting device is arranged on a core machine of the gas turbine engine and collects redundant high-energy working medium generated by the high-energy working medium collecting device. The output end of the high-energy working medium collecting device is connected with the high-efficiency working medium transmission device, and the high-energy working medium is transmitted to the distributed propellers through the high-efficiency working medium transmission device. In the distributed propeller, the energy of high-energy working medium is used for driving the propeller to generate thrust.
The gas turbine engine core is a single-rotor or double-rotor gas turbine jet engine; high-energy working medium is generated through the Brayton cycle in the device. The high-energy working medium is compressed air led out from the core compressor, and can be reheated at the output end of the turbine, so that the internal energy of the working medium is further improved.
The high-energy working medium collecting device comprises at least one gas collecting device, at least one set of transmission pipeline and at least one tail gas heat regenerative device. At least one set of gas collecting device is arranged on each gas turbine engine core. The input end of the gas collecting device is communicated with the output end of the compression part of the core machine of the gas turbine engine, and the redundant high-energy working medium generated by the core machine of the gas turbine engine is collected. The output end of the gas collecting device is communicated with the input end of the transmission pipeline, and the working medium is transmitted to at least one tail gas heat recovery device through the transmission pipeline. The output end of the transmission pipeline is communicated with the input end of the tail gas heat recovery device, and the compressed air exchanges heat with the tail gas in the tail gas heat recovery device. The output end of the tail gas heat regenerative device is communicated with the input end of the high-efficiency working medium conveying device.
The high-efficiency working medium transmission device is a high-thermal-resistance and low-resistance working medium transmission pipeline and is used for transmitting high-energy working media to the distributed propellers. The high-efficiency working medium transmission device comprises at least one transmission main pipe and a plurality of transmission branch pipes. The input end of the transmission main pipe (namely the input end of the high-efficiency working medium transmission device) is connected with the output end of the working medium collecting device, and the output end of the transmission main pipe is a multi-path transmission branch pipe connected in parallel. The input end of the transmission branch pipe is connected with the transmission main pipe; the output end of the transmission branch pipe is connected with the distributed propeller. Each distributed propeller corresponds to at least one transmission branch pipe. The transmission main pipe and the transmission branch pipes are made of high-temperature-resistant and high-pressure-resistant materials, and flow guide devices are arranged at the corners inside the pipelines to reduce flow loss.
The function of the distributed thruster is to convert the energy of the high-energy working medium into the propulsive work of the aircraft, and the realization mode of the distributed thruster comprises but is not limited to a gear transmission fan system driven by a turbine, such as a gear transmission propeller system driven by the turbine. For a gear transmission fan system driven by a propeller turbine, high-energy working medium sequentially passes through a transmission main pipe and a transmission branch pipe of a high-efficiency working medium transmission device to enter the propeller turbine and drive the propeller turbine to operate. The impeller turbine drives the fan to operate through the gear reducer, and thrust is generated.
Compared with the prior art, the invention can further improve the efficiency of the power system of the existing conveyer, reduce carbon emission, save energy and protect environment, and simultaneously improve the realizability of the distributed propulsion system. The specific analysis is as follows:
At present, the bypass ratio of the civil turbofan engine with a large bypass ratio is close to 10, and the limit which can be reached in the future is 10-15. Aiming at the above example, a pneumatic thermodynamic calculation model is established, and the influence of main design parameters on the bypass ratio and the oil consumption rate of the propulsion system is analyzed, as shown in fig. 10 to 13. The result of the pneumatic-thermal calculation of the distributed propulsion system based on the working medium shows that the bypass ratio of the distributed propulsion system is expected to reach more than 20-25 under the prior art, and the oil consumption rate is reduced by 40% -50% compared with the oil consumption rate of the most common CFM56 engine (the bypass ratio is about 6.0, and the takeoff oil consumption rate is about 0.37 kg/kgf/h).
Drawings
FIG. 1 is a schematic structural view of the present invention;
FIG. 2 is a schematic structural view of embodiment 1;
FIG. 3 is a schematic configuration diagram of a gas turbine engine core machine in embodiment 1;
FIG. 4 is a schematic configuration diagram of a gas turbine engine core machine in embodiment 2;
FIG. 5 is a schematic view of the gas turbine engine core compressor discharge ring configuration; wherein 5a is an isometric view of the gas collecting ring, 5b is a front view, and 5c is a partial enlarged view of 5 b;
FIG. 6 is a schematic diagram of a U-shaped tube bundle heat regenerator; wherein 6a is a front view, 6b is a side view, and 6c is a top view;
fig. 7 is a schematic structural view of a distributed propeller in embodiment 1; wherein 7a is a front view, 7b is a side view, and 7c is a top view;
fig. 8 is a schematic structural view of a distributed propeller in embodiment 3; wherein 8a is a front view, 8b is a side view, and 8c is a top view;
FIG. 9 is a layout of a distributed propulsion system on an aircraft; wherein 9a is a front view and 9b is a top view;
FIG. 10 is an illustration of the effect of core boost ratio on bypass ratio and fuel consumption; wherein 10a is the influence of the boost ratio on the bypass ratio, 10b is the influence of the boost ratio on the fuel consumption; in fig. 10, PRC means a pressure increase ratio, BPR means a bypass ratio, and sfc means a fuel consumption rate.
FIG. 11 is a core turbine front total temperatureInfluence on bypass ratio and fuel consumption; wherein 11a is the influence of the front total temperature of the core turbine on the bypass ratio, and 11b is the influence of the front total temperature of the core turbine on the oil consumption rate; t in FIG. 11 4 The temperature of the core engine before turbine is indicated, BPR is a bypass ratio, and sfc is the oil consumption rate.
FIG. 12 is the effect of working medium regenerative temperature on bypass ratio and fuel consumption; wherein 12a is the influence of the regenerative temperature of the working medium on the bypass ratio, and 12b is the influence of the regenerative temperature of the working medium on the oil consumption rate; in FIG. 12, T 42 The heat regeneration temperature of the working medium is indicated, the BPR is the bypass ratio, and the sfc is the oil consumption rate.
FIG. 13 is an illustration of the effect of total pressure recovery on bypass ratio and fuel consumption of a working fluid transfer device; wherein 13a is the influence of the total pressure recovery coefficient of the working medium transmission device on the bypass ratio, and 13b is the influence of the total pressure recovery coefficient of the working medium transmission device on the oil consumption rate; in fig. 13, σ denotes a total pressure recovery coefficient, BPR denotes a bypass ratio, and sfc denotes a fuel consumption rate. In the figure:
1. an internal support of the distributed thruster; 2. a high-energy working medium collecting device; 3. a gas collection device; 4. a transmission pipeline; 5. a tail gas heat regenerative device; 6. a high pressure compressor; 7. a high-efficiency working medium transmission device; 8. a transmission main pipe; 9. a high pressure turbine; 10. conveying branch pipes; 11. a high pressure turbine shaft; 12. a distributed thruster; 13. a flow guide device; 14. a gas turbine engine core; 15. a core machine compressor; 16. a core engine combustion chamber; 17. a core turbine; 18. a core machine nozzle; 19. a propeller turbine; 20. a gear reducer; 21. a fan; 22. a propeller; 23. the propeller is externally connected with a culvert spray pipe; 24. a distributed propulsion system; 25. a wing-body fusion aircraft; 26. a propulsion pod; 27. a distributed thruster support plate; 28. a propeller culvert spray pipe; 29. a casing in the distributed thruster; 30. an air duct; 31. a compressor outer case; 32. a turbine rear case; 33. a gas collecting ring; 34.U-shaped tube bundle heat regenerator; 35. an input end of a heat regenerator; 36. an output end of the heat regenerator; 37.U-shaped heat exchange tubes; 38.U-shaped heat exchange tube output end; 39. a hot gas input; 40. a hot gas output end; 41. a cold incoming flow header; 42. a hot return manifold; 43. a hot air duct; the input end of the U-shaped heat exchange tube; 45. a propeller turbine volute; 46. a propeller turbine shaft; 47. a ducted fan case; 48. a gas-guide tube regulating valve; 49 bearing; 50. a low pressure compressor; 51. a low pressure turbine; 52. low-pressure turbine shaft
Detailed Description
Example 1
The embodiment is a distributed propulsion system based on working medium transmission.
The embodiment comprises a turbine engine core machine 14, a high-efficiency working medium transmission device 7, a high-energy working medium collection device 2 and a distributed propeller 12. The input end of the gas collecting device 3 of the high-energy working medium collecting device 2 is communicated with the output end of a core machine compressor 15 of a core machine of a turbine engine; the output end of the high-energy working medium collecting device 2 is communicated with the inlet of the high-efficiency working medium transmission device 7. The output end of the transmission branch pipe 10 in the high-efficiency working medium transmission device 7 is respectively communicated with the input end of the propeller turbine volute 45 of each distributed propeller 12; the input end of the transmission branch pipe 10 is communicated with a transmission main pipe 8 in the high-efficiency working medium transmission device 7. In this embodiment, there are six distributed thrusters 12; six distributed propellers 12 are equally distributed on either side of the turbine engine core 14. In this embodiment, the turbine engine core 14 is a single spool gas turbine jet engine.
The gas collecting device 3 comprises four gas guide tubes 30, a gas collecting ring 33 and four gas guide tube regulating valves 48. The gas collecting ring 33 is an annular shell, four gas guide tube connecting holes are uniformly distributed on the surface of the same side of the shell of the gas collecting ring 33, and the input ends of the four gas guide tubes 30 are respectively arranged on the gas guide tube connecting holes. The four airway regulating valves 48 are respectively mounted at the output ends of the airway tubes 30. The surface of the casing of the gas collecting ring 33 is arc-shaped, and the opening of the casing of the arc-shaped gas collecting ring 33 is positioned at the inner side of the casing of the gas collecting ring 33, so that the cross section of the casing is U-shaped. The compressor outer casing 31 is divided into two sections; two side walls of the opening of the gas collecting ring 33 are respectively and fixedly connected with the compressor casing 31 which is divided into two sections, so that the gas collecting ring 33 and the compressor casing 31 jointly form an outer ring channel of the gas flow at the outlet of the compressor. The core compressor 15 on the gas turbine engine core 14 has two annular outlets, namely an inner annular outlet and an outer annular outlet; the outer annular outlet is communicated with the input end of the air flow outer annular channel at the outlet of the compressor; the inner annular outlet communicates with the input of the core combustor 16.
During assembly, the output ends of the four gas guide pipes 30 on the casing of the gas collecting ring 33 are respectively communicated with the input ends of the four tail gas heat recovery devices 5 distributed in the rear casing 32 of the turbine, and the gas flow of the outer ring channel of the gas flow at the outlet of the compressor is transmitted into each tail gas heat recovery device 5 through each gas guide pipe 30. The 4 tail gas regenerative devices are divided into two groups, and the output ends of the tail gas regenerative devices in each group are respectively communicated with the transmission main pipes 8 positioned at two sides of the turbine engine core 14 through hot air guide pipes 43.
The exhaust gas heat recovery device 5 is circumferentially and uniformly installed inside the turbine rear casing 32, and an included angle α =0 ° to 90 ° between an axis of the exhaust gas heat recovery device and an axis of the turbine engine core 14.
In this embodiment, the tail gas heat recovery device 5 is a U-shaped tube bundle heat recovery device 34. The regenerator inlet 35 is connected to a cold incoming flow manifold 41. The multiple U-shaped heat exchange tubes 37 are connected in parallel on the side wall of the cold incoming flow header 41. The air flow enters the U-shaped heat exchange tube 37 through the U-shaped heat exchange tube input 44 on the wall of the incoming flow manifold 41. The hot gas of the U-shaped heat exchange tube 37 flows through the gap between the outer sides of the U-shaped heat exchange tube 37, and exchanges heat with the compressed air inside the U-shaped heat exchange tube 37 through the tube wall. The heat exchanged hot air is collected from the U-shaped heat exchange tube output 38 to a heat return manifold 42 and passes through the heat return manifold 42 to the regenerator output 36. Each regenerator output 36 is connected to a respective hot air duct 43. The output end of the hot air conduit 43 is connected with the transmission manifold 8 of the high-efficiency working medium transmission device 7. The high-temperature and high-pressure air after heat exchange enters the transmission header 8 through the hot air conduit 43.
Two of the transfer manifolds 8 are located on either side of the turbine engine core 14. The output end of the group of U-shaped tube bundle heat regenerators 34 is connected to the hot air duct 43 and is respectively connected to the input end of each transmission main pipe through a three-way joint.
Three airflow output ends are distributed on each transmission main pipe 8, and each airflow output end is connected with a transmission branch pipe 10. The output end of the transmission branched pipe 10 is respectively communicated with the input end of the propeller turbine volute 45 of each distributed propeller 12.
The distributed propeller 12 uses a propeller turbine 19 to drive a fan 21.
When the distributed impeller is a ducted fan driven by a propeller turbine, the ducted fan includes a fan 21, a ducted fan casing 47, a gear reducer 20, a propeller turbine shaft 46, a propeller turbine volute 45, a propeller turbine 19, a distributed impeller inner casing 29, and a distributed impeller strut 27. The number of the distributed propeller support plates 27 is 4, and the distributed propeller support plates are uniformly distributed between the ducted fan casing 47 and the distributed propeller inner casing 29, so that one end of each distributed propeller support plate 27 is fixed on the inner surface of the ducted fan casing 47, and the other end of each distributed propeller support plate is fixed on the outer surface of the distributed propeller inner casing 29. The propeller turbine shaft 46 is positioned in the inner casing 29 of the distributed propeller, and both ends of the propeller turbine shaft are arranged on the inner bracket 1 of the distributed propeller through a bearing 49; the bracket distributed propeller inner bracket 1 is fixed on the inner surface of a distributed propeller inner casing 29; the centerline of the propeller turbine shaft 46 coincides with the centerline of the distributed propeller inner casing 29.
The gear reducer 20 is sleeved at the front end of the propeller turbine shaft 46; the propeller turbine 19 is sleeved on the rear end of the propeller turbine shaft 46.
The gear reducer 20 is a planetary gear reducer; a sun gear of the planetary gear reducer is sleeved at the front end of the propeller turbine shaft 46; the fan 21 is fitted over the outer gear ring of the planetary gear reducer.
The input end of the transmission branched pipe 10 passes through the through hole on the distributed propeller support plate 27 and is connected with the input end of the propeller turbine volute 45. The propeller turbine volute 45 is annular and is wrapped around the propeller turbine 19, and the output ends of the propeller turbine volute 45 are uniformly distributed on the inner side of the propeller turbine volute 45 and are connected with the input end of the centrifugal propeller turbine 19.
The propeller turbine 19 converts the internal energy of the high-temperature and high-pressure air into shaft work, and drives the fan 21 to operate through the propeller turbine shaft 46 and the gear reducer 20. An optimum matching of the rotational speed of the fan 21 to the propeller turbine 19 is achieved by means of the gear reduction 20. The fan 21 draws in ducted fan case 47 and compresses the air. The compressed air then expands within the propeller bypass nozzle 23 accelerating the discharge to produce the main thrust.
The working mode of the embodiment is as follows:
the gas turbine engine core 14 utilizes the brayton cycle occurring therein to produce high energy working fluid. The core compressor 15 draws in and compresses incoming air. The compressed air is divided into two streams at the output end of the core machine compressor 15, one stream of air flow enters the air collecting ring 33 from the input end of the air collecting ring 33, and the other stream of air flow enters the core machine combustion chamber 16 to be mixed and combusted with fuel, so that high-temperature and high-pressure fuel gas is formed. The combustion gases expand in the core turbine 17 to produce work and drive the core compressor 15 into operation. The high-temperature and high-pressure gas behind the core turbine 17 continuously passes through the tail gas heat recovery device 5 to exchange heat with the compressed air. The combustion gases are discharged through the core nozzle 18, producing a small amount of thrust.
Air collected by the air collector ring 33 is delivered through air duct 30 to the inlet of the U-tube bundle regenerator 34 and into the cold incoming flow manifold 41. The air flow enters the U-shaped heat exchange tube 37 through the U-shaped heat exchange tube input 44 on the wall of the incoming flow manifold 41. The hot gas flows through the gap between the outer sides of the U-shaped heat exchange tubes 37 and exchanges heat with the compressed air inside the U-shaped heat exchange tubes 37 through the tube walls. The heat exchanged hot air is collected from the U-shaped heat exchange tube output 38 to a heat return manifold 42 and passes through the heat return manifold 42 to the regenerator output 36. Each regenerator output 36 is connected to a respective hot air duct 43. The hot air duct 43 introduces the high-temperature and high-pressure air after heat exchange into the transmission header 8 through the hot air duct 43.
The high-temperature and high-pressure compressed air is regulated by the air duct regulating valve 48 from the transmission main pipe 8 to gradually enter the transmission branch pipe 10 and finally reach the plurality of distributed propellers 12.
The high temperature and high pressure air enters the propeller turbine volute 45 through the transfer manifold 10. The impeller turbine volute 45 directs high temperature, high pressure air into the impeller turbine 19. The propeller turbine 19 converts the expansion work of the high-temperature and high-pressure air into shaft work, and drives the fan 21 to operate through the propeller turbine shaft 46 and the gear reducer 20. An optimum matching of the rotational speed of the fan 21 to the propeller turbine 19 is achieved by means of the gear reduction 20. The fan 21 draws in ducted fan case 47 and compresses the air. The compressed air then expands within the propeller bypass nozzle 23 accelerating the discharge to produce the main thrust.
Fig. 9 is a layout of an example of the distributed propulsion system 24 on a wing-body fusion aircraft 25. The distributed propulsion system 24 is located in a propulsion nacelle 26 on the upper aft side of the wing-body fusion vehicle 25. A single gas turbine engine core 14 is located on the fuselage centerline of a wing- body fusion aircraft 25, and 6 geared fan systems 22 are symmetrically disposed on the left and right sides of the gas turbine engine core 14. Such an arrangement allows the engine to utilize the low speed boundary layer flowing over the surface of the fuselage fusion vehicle 25, improving propulsion efficiency.
Example 2
The embodiment is a distributed propulsion system based on working medium transmission.
The embodiment comprises a turbine engine core machine 14, a high-efficiency working medium transmission device 7, a high-energy working medium acquisition device 2 and a distributed propeller 12. The input end of the gas collecting device 3 of the high-energy working medium collecting device 2 is communicated with the output end of a high-pressure compressor 6 of a core machine of a turbine engine; and the output end of the high-energy working medium collecting device 2 is communicated with the inlet of the high-efficiency working medium transmission device. The output end of the transmission branch pipe 10 in the high-efficiency working medium transmission device 7 is respectively communicated with the input end of the propeller turbine volute 45 of each distributed propeller 12; the input end of the transmission branch pipe 10 is communicated with a transmission main pipe 8 in the high-efficiency working medium transmission device 7. In this embodiment, there are six distributed thrusters 12; six distributed propellers 12 are equally distributed on either side of the turbine engine core 14. In this embodiment, the turbine engine core 14 is a dual rotor gas turbine jet engine.
Two side walls of the opening of the gas collecting ring 33 are respectively and fixedly connected with the compressor outer casing 31 which is divided into two sections, so that the gas collecting ring 33 and the compressor outer casing 31 jointly form an outer ring channel of the outlet airflow of the high-pressure compressor. The high-pressure compressor 6 on the gas turbine engine core 14 is provided with two annular outlets, namely an inner annular outlet and an outer annular outlet; the outer annular outlet is communicated with the input end of the outer annular channel of the airflow at the outlet of the high-pressure compressor; the inner annular outlet communicates with the input of the core combustor 16.
During assembly, the output ends of the four gas guide pipes 30 on the casing of the gas collecting ring 33 are respectively communicated with the input ends of the four tail gas heat recovery devices 5 distributed in the rear casing 32 of the turbine, and the airflow of the outer annular channel of the airflow at the outlet of the high-pressure compressor is transmitted into each tail gas heat recovery device 5 through each gas guide pipe 30. The 4 tail gas heat recovery devices are divided into two groups, and the output ends of the tail gas heat recovery devices of each group are respectively communicated with the transmission main pipes 8 positioned at two sides of the turbine engine core 14 through hot air guide pipes 43.
Example 3
The embodiment is a distributed propulsion system based on working medium transmission.
The embodiment comprises a turbine engine core machine 14, a high-efficiency working medium transmission device 7, a high-energy working medium collection device 2 and a distributed propeller 12. The input end of the gas collecting device 3 of the high-energy working medium collecting device 2 is communicated with the output end of a high-pressure compressor 6 of a core machine of a turbine engine; and the output end of the high-energy working medium collecting device 2 is communicated with the inlet of the high-efficiency working medium transmission device. The output end of the transmission branch pipe 10 in the high-efficiency working medium transmission device 7 is respectively communicated with the input end of the propeller turbine volute 45 of each distributed propeller 12; the input end of the transmission branch pipe 10 is communicated with the transmission main pipe 8 in the high-efficiency working medium transmission device 7. In this embodiment, there are six distributed thrusters 12; six distributed propellers 12 are evenly distributed on either side of the turbine engine core 14.
Three airflow output ends are distributed on each transmission main pipe 8, and each airflow output end is connected with a transmission branch pipe 10. The output end of the transmission branched pipe 10 is respectively communicated with the input end of the propeller turbine volute 45 of each distributed propeller 12.
The distributed thruster 12 drives a propeller 22 with a thruster turbine 19.
The distributed thruster comprises a propeller 22, a gear reducer 20, a thruster turbine shaft 46, a thruster turbine volute 45, a thruster turbine 19 and a distributed thruster inner casing 29. The propeller turbine shaft 46 is positioned in the inner casing 29 of the distributed propeller, and both ends of the propeller turbine shaft are arranged on the inner support 1 of the distributed propeller through bearings 49; the bracket distributed propeller inner bracket 1 is fixed on the inner surface of a distributed propeller inner casing 29; the centerline of the propeller turbine shaft 46 coincides with the centerline of the distributed propeller inner casing 29.
The gear reducer 20 is sleeved at the front end of the propeller turbine shaft 46; the propeller turbine 19 is sleeved on the rear end of the propeller turbine shaft 46.
The gear reducer 20 is a planetary gear reducer; a sun gear of the planetary gear reducer is sleeved at the front end of the propeller turbine shaft 46; the propeller 22 is fitted over the outer gear ring of the planetary gear reducer.
The output end of the transmission branched pipe 10 is connected with the input end of the propeller turbine volute 45. The propeller turbine volute 45 is annular and is wrapped around the propeller turbine 19, and the output ends of the propeller turbine volute 45 are uniformly distributed on the inner side of the propeller turbine volute 45 and are connected with the input end of the centrifugal propeller turbine 19.
The propeller turbine 19 converts the internal energy of the high-temperature and high-pressure air into shaft work, and drives the propeller 22 to rotate through the propeller turbine shaft 46 and the gear reducer 20. An optimum matching of the rotational speed of the propeller 22 to the propeller turbine 19 is achieved by means of the gear reduction 20. The propeller 22 pushes the airflow backwards, creating a pulling force.
Claims (9)
1. A distributed propulsion system is characterized by comprising a turbine engine core machine, a high-efficiency working medium transmission device, a high-energy working medium collection device and a distributed propeller; the input end of the gas collection device of the high-energy working medium collection device is communicated with the output end of a core machine compressor of a turbine engine core machine; the output end of the high-energy working medium collecting device is communicated with the inlet of the high-efficiency working medium transmission device; the output end of a transmission branch pipe in the high-efficiency working medium transmission device is respectively communicated with the input end of a propeller turbine volute of each distributed propeller; the input end of the transmission branch pipe is communicated with a transmission main pipe in the high-efficiency working medium transmission device; 6 distributed propellers are evenly distributed on two sides of the turbine engine core.
2. The distributed propulsion system as claimed in claim 1, wherein the turbine engine core is a single spool gas turbine jet engine or a twin spool turbine jet engine; the distributed propellers are ducted fans driven by propeller turbines or propellers driven by the propeller turbines.
3. The distributed propulsion system of claim 1, wherein the air collection device includes four air ducts, an air collection ring, and four air duct adjustment valves; the surface of the same side of the gas collecting ring shell is uniformly provided with four gas guide pipe connecting holes, and the input ends of the four gas guide pipes are respectively arranged on the gas guide pipe connecting holes; the four air duct regulating valves are respectively arranged at the output ends of the air ducts; the surface of the gas collecting ring shell is arc-shaped, and the opening of the arc-shaped gas collecting ring shell is positioned on the inner side of the gas collecting ring shell, so that the cross section of the arc-shaped gas collecting ring shell is U-shaped; cutting the compressor outer casing into two sections; two side walls of the opening of the gas collecting ring are respectively fixedly connected with a compressor outer casing which is divided into two sections, so that the gas collecting ring and the compressor outer casing jointly form a compressor outlet airflow outer ring channel; the core compressor on the turbine engine core machine is provided with two annular outlets, namely an inner annular outlet and an outer annular outlet; the outer annular outlet is communicated with the input end of the air flow outer annular channel at the outlet of the compressor; the inner annular outlet is communicated with the input end of the core machine combustion chamber;
during assembly, the output ends of four gas guide pipes on the gas collecting ring shell are respectively communicated with the input ends of four tail gas heat recovery devices distributed in a rear casing of the turbine, and the gas flow of an outer ring channel of the gas flow at the outlet of the gas compressor is transmitted into each tail gas heat recovery device through each gas guide pipe; the 4 tail gas regenerative devices are divided into two groups, and the output ends of the tail gas regenerative devices of each group are respectively communicated with the transmission main pipes positioned at two sides of the turbine engine core machine through hot air guide pipes.
4. The distributed propulsion system as claimed in claim 3, wherein the exhaust gas heat recovery device is circumferentially and uniformly mounted inside the turbine rear casing, and an included angle α =0 ° to 90 ° between an axis of the exhaust gas heat recovery device and an axis of the turbine engine core.
5. The distributed propulsion system as claimed in claim 1, wherein there are two of the transfer manifolds, one on each side of the turbine engine core; the output ends of the U-shaped tube bundle heat regenerators are communicated with hot air guide tubes and are respectively communicated with the input ends of the transmission main pipes through three-way joints; three airflow output ends are distributed on each transmission main pipe, and each airflow output end is connected with a transmission branch pipe; and the output end of the transmission branch pipe is respectively communicated with the input end of the propeller turbine volute of each distributed propeller.
6. The distributed propulsion system as claimed in claim 1 wherein the input end of the transport veins passes through the through holes in the distributed thruster support plate and is connected to the input end of the thruster turbine volute; the propeller turbine volute is annular and is wrapped around the propeller turbine, and the output end of the propeller turbine volute is uniformly distributed on the inner side of the propeller turbine volute and is connected with the input end of the centrifugal propeller turbine.
7. The distributed propulsion system of claim 2, wherein when the distributed propulsor is a propulsor turbine driven ducted fan, the ducted fan includes a fan, a ducted fan case, a gear reducer, a propulsor turbine shaft, a propulsor turbine volute, a propulsor turbine, a distributed propulsor inner case, and a distributed propulsor strut; the 4 distributed propeller support plates are uniformly distributed between the ducted fan casing and the distributed propeller inner casing, so that one end of each distributed propeller support plate is fixed on the inner surface of the ducted fan casing, and the other end of each distributed propeller support plate is fixed on the outer surface of the distributed propeller inner casing; the propeller turbine shaft is positioned in the inner casing of the distributed propeller, and two ends of the propeller turbine shaft are both arranged on the inner support of the distributed propeller through bearings; the inner support of the distributed propeller is fixed on the inner surface of the inner casing of the distributed propeller; the center line of the propeller turbine shaft is superposed with the center line of the casing in the distributed propeller.
8. The distributed propulsion system as claimed in claim 2 wherein when the turbine engine core employs a dual rotor turbojet engine, the output of the high pressure compressor of the turbine engine core communicates with the input of the gas collection device of the high energy working medium collection device; and the output end of the high-energy working medium collecting device is communicated with the inlet of the high-efficiency working medium conveying device.
9. The distributed propulsion system as claimed in claim 2 wherein when the distributed propulsion means is a propeller driven by a propeller turbine, the propeller pushes the airflow backwards, creating a drag force.
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| Application Number | Priority Date | Filing Date | Title |
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| CN201710008820.9A CN106742075B (en) | 2017-01-06 | 2017-01-06 | Distributed propulsion system |
| EP17879560.5A EP3566952B1 (en) | 2017-01-06 | 2017-12-27 | Distributed propulsion system |
| DE112017000168.4T DE112017000168T5 (en) | 2017-01-06 | 2017-12-27 | Distribution drive system |
| PCT/CN2017/118908 WO2018103762A1 (en) | 2017-01-06 | 2017-12-27 | Distributed propulsion system |
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| CN201710008820.9A CN106742075B (en) | 2017-01-06 | 2017-01-06 | Distributed propulsion system |
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| DE112017000168T5 (en) * | 2017-01-06 | 2018-08-16 | Northwestern Polytechnical University | Distribution drive system |
| CN107745818B (en) * | 2017-10-10 | 2020-05-19 | 中国人民解放军国防科技大学 | Aircraft propulsion system and vertical take-off and landing aircraft with same |
| CN110005544A (en) * | 2019-05-12 | 2019-07-12 | 西北工业大学 | From driving by-pass air duct annular flabellum compression set |
| CN110588878B (en) * | 2019-09-20 | 2021-06-22 | 辽宁壮龙无人机科技有限公司 | Manufacturing method of propeller and propeller |
| CN111332464B (en) * | 2020-03-02 | 2021-07-20 | 中国空气动力研究与发展中心高速空气动力研究所 | Distributed propulsion flying wing aircraft |
| CN111661344B (en) * | 2020-07-13 | 2021-09-24 | 中国航空发动机研究院 | Wing-body integrated aircraft propulsion system |
| CN112443423B (en) * | 2020-11-24 | 2022-04-05 | 南京航空航天大学 | Jet propulsion power system of air-driven ducted fan |
| CN114439646B (en) * | 2022-01-27 | 2022-12-06 | 西北工业大学 | Air turbine rocket stamping combined propulsion system |
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